The present invention relates to a turbine engine. In particular, the invention relates to liner cooling for combustor for a gas turbine engine.
A turbine engine ignites compressed air and fuel in a combustion chamber, or combustor, to create a flow of hot combustion gases to drive multiple stages of turbine blades. The turbine blades extract energy from the flow of hot combustion gases to drive a rotor. The turbine rotor drives a fan to provide thrust and drives compressor to provide a flow of compressed air. Vanes interspersed between the multiple stages of turbine blades align the flow of hot combustion gases for an efficient attack angle on the turbine blades.
There is a desire to improve the fuel efficiency, or thrust specific fuel consumption (TSFC), of turbine engines. TSFC is a measure of the fuel consumed per unit of thrust produced by an engine. Fuel efficiency may be improved by increasing the combustion temperature and pressure under which the engine operates. However, under such conditions, undesirable combustion byproducts (e.g. nitrogen oxides (NOx)) may form at an increased rate. In addition, the higher temperatures may require additional cooling air to protect engine components. A source of cooling air is typically taken from a flow of compressed air produced upstream of the turbine stages. Energy expended on compressing air used for cooling engine components is not available to produce thrust. Improvements in the efficient use of compressed air for cooling engine components can improve the overall efficiency of the turbine engine.